Spacecraft Design March 2, 2009

1

Xenia Spacecraft StudyXenia Spacecraft Study

Spacecraft DesignSpacecraft Design

March 2, 2009March 2, 2009

2

OutlineOutline

• Study Overview• Spacecraft Team Members• Animation• Overall Ground Rules and Assumptions (GR&A)• Mission Analysis • Configuration • Mass Properties• Guidance, Navigation, and Control (GN&C)• Avionics• Power• Thermal• Propulsion• Structures• Conclusions

3

Study OverviewStudy Overview

• Goal– Perform a mission concept study for the proposed Xenia mission

• Responsibilities

• Spacecraft: ED04 – Avionics / GN&C– Communications– Electrical Power– Trajectory / Mission Analysis– Propulsion– Science Instruments Integration– Launch Stack Shroud Integration– Animation / Modeling

• Science: VP62– Science Instruments

Definition– Science Instruments Design– Mission requirements

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Spacecraft Team MembersSpacecraft Team Members

Chryssa Kouveliotou Xenia PI MSFC-VP62

Les Johnson Study Manager and Lead MSFC-ED04

Randy Hopkins Technical Lead / Mission Analysis MSFC-ED04

Mike Baysinger Spacecraft Configuration MSFC-ED04 / Qualis

P.J. Benfield Propulsion U. Of Alabama Huntsville

Pete Capizzo Avionics / GN&C MSFC-ED04 / Raytheon

Tracie Crane Mass Properties MSFC-ED04 / Qualis

Leo Fabisinski Power MSFC-ED04 / ISSI

Linda Hornsby Thermal MSFC-ED04 / JTI

David Jones Structures MSFC-ED04

Dauphne Maples Mass Properties / GR&A MSFC-ED04 / Qualis

Janie Miernik Structures MSFC-ED04 / ERC

Tom Percy Mission Analysis SAIC

Kevin Thompson Animation MSFC-ED04 / Jacobs

Matt Turner Propulsion U. Of Alabama Huntsville

5

AnimationAnimation

• The Spacecraft Engineering team has created an animation that depicts the science mission of the Xenia spacecraft.

• Link to the animation:– http://sms.msfc.nasa.gov/xenia/

6

• Additional GR&A are contained in each discipline section.

• Preferred Launch Vehicle is the Falcon 9, launched from Omelek (Kwajalein).

• Target orbit is 600km circular, 5-degree inclination (or less).

• Target spacecraft lifetime = 5 years.

• Target orbit lifetime = 10 years.

• Science instruments designed by VP62.– Instrument parameters (power, mass, etc.) provided by VP62.

Overall Ground Rules and Overall Ground Rules and Assumptions (GR&A)Assumptions (GR&A)

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Mission AnalysisMission Analysis

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Mission Analysis: GR&AMission Analysis: GR&A

• Launch Vehicle Performance– Target orbit is 600km circular, inclination no greater than 5 degrees

Avoid the South Atlantic Anomaly

– Preferred launch vehicle is Falcon 9 Launched from Omelek (Kwajalein) Payload adapter mass has been subtracted from the payload performance quotes

• Orbital Lifetime– Reentry interface defined as 400000ft altitude (122 km)– Initial Circular orbit altitude = 600 km– Target lifetime is 10 years– Start dates are July 1, 2012, and July 1, 2018, in order to capture the effect

of the solar maximum– Use the orbital lifetime tool included in Satellite Toolkit (STK)

9

Mission Analysis: Launch VehiclesMission Analysis: Launch Vehicles

Falcon 9 Vega Atlas V 401 Delta II Heavy (7920H-10)*

Launch Site Omelek (Kwajalein)

CCAFS Kourou CCAFS CCAFS

Source: NASA LSP [2] NASA LSP [2] Vega User’s Manual [1]

NASA LSP [2] NASA LSP [2]

600 km @ 5 deg

7000 1700 2050 4395 895

600 km @ 10 deg

Not requested

TBD [3] 2040 5815 1440

600 km @ 15 deg

Not requested

TBD [3] Not requested

Not requested

Not requested

500 km @ 5 deg

Not requested

Not requested

2120 4390 885

500 km @ 10 deg

Not requested

Not requested

2110 5820 1435

* Also known as the 2920H-10 and 2925H-10.[1] Interpolated results from performance plots. The mass of the 60kg type 937 payload

adapter has been subtracted.[2] The Falcon, Atlas, and Delta II guides do not include performance estimates for these low

inclinations.[3] Data is pending, but was not available at the time of this briefing.

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Mission Analysis: Orbital LifetimeMission Analysis: Orbital Lifetime

Orbital Lifetime for two spacecraft masses, 2012 Launch Date

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Date

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tude

(km

)

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2100 kg

10 y

ears

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Mission Analysis: Orbital LifetimeMission Analysis: Orbital Lifetime

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tude

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2100 kg10

yea

rs

Orbital Lifetime for two spacecraft masses, 2018 Launch Date

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Mission Analysis: DeMission Analysis: De--orbitorbit

– Delta-V = 163 m/s for a reentry flight path angle of -1.75 degrees

Impulsive Delta-V: 161.3 m/s Gravity Loss: 1.7 m/s (assuming worst case T/W = 0.025) Margin: 0 m/s (assumptions are already conservative)

– Perigee altitude = 34.6 km

Ranges from 65.7 km to 8.25 km for the acceptable range of reentry flight path angles

– Gravity Loss is insignificant for T/W > 0.025

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Mission Analysis: ConclusionsMission Analysis: Conclusions

• Launch Vehicle– Falcon 9 has large mass margin if launched from Omelek.– Atlas V 401 launched from CCAFS provides large mass margin.– Delta II Heavy has insufficient payload mass.– Vega has insufficient payload mass; envelope too small.

• Orbital Lifetime– Based on the calculations, no periodic orbit boost will be required.

• De-orbit– Need a propulsion system which can supply a total delta-v of 163 m/s

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Spacecraft ConfigurationSpacecraft Configuration

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Configuration: Falcon ShroudConfiguration: Falcon Shroud

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ConfigurationConfiguration

2.7

4

Ø2.6

2.5

17


bulkhead

R42 tilted to go thru c.g.

WFI boxes

WFS boxes

bulkheadscope support



scope support

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bulkhead

R42 tilted to go thru c.g.

WFI

WFS




scope support

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Configuration: DeConfiguration: De--orbit systemorbit system

200 lb R42thruster

NTO/MMH tanks

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Configuration: Avionics BoxesConfiguration: Avionics Boxes

Data Recorder

Batt ChargersFlight Computer

Battery

WFM Boxes

MGA

ARU

Battery

WFSWFI

WFM

WFIWFS

WFM

CMGs

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Configuration: CMG locationsConfiguration: CMG locations

CMGs 90 deg apart

Ixx=3092 kg*m2

Iyy=1900 kg*m2

Izz=3317 kg*m2X

Z

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Configuration: Array deploymentConfiguration: Array deployment

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◙.15

1.4 2.7

4

2.5

10.5

c.g.

3

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Mass PropertiesMass Properties

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Mass Properties: Mass Properties: GR&A and ResultsGR&A and Results

• GR&A– Growth Allowance

Spacecraft: 30% Science Instruments: Obtained through VP62 Science Team

• Results– Mass Total 30% Margin: 2753.83 kg– Mass Total 20% Margin: 2640.56 kg– Mass Total 0% Margin: 2414.03 kg

Note: Above margins do not apply to science instruments.

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Mass Properties: ResultsMass Properties: Results

WBS Element - Descent Stage Qty Unit Mass (kg) Total Mass (kg)1.0 489.002.0 15.503.0 169.524.0 425.945.0 32.706.0 339.80

6.1 30% 146.706.2 30% 4.656.3 30% 50.866.4 30% 127.786.5 Thermal 30% 9.81

1472.477.0 6.108.0 1138.00

8.1 1 575.00 575.008.2 1 384.00 384.008.3 1 144.00 144.008.4 Instrument Cabling 1 35.00 35.00

Inert Mass 1144.10Total Less Propellant 2616.57

9.0 137.26Gross Mass 2753.83

WFM

Propellant

WFSWFI

Thermal ControlGrowth

Dry Mass

Power

StructurePropulsion

Cargo/PayloadNon-Cargo

Avionics/Control

StructurePropulsionPowerAvionics/Control

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GN&C: GR&AGN&C: GR&A

• Operational Pointing/viewing coverage– 360deg (entire sky), with 45deg sun avoidance – no earth or moon avoidance required

• Fast slew requirements– autonomous slew of 60deg /60sec to detected target– At least once in a 24 hr period

• Slow slew requirements– Up to 5 slow slews per orbit, 100 deg per slew

• Pointing accuracy: – after fast slew - within 2 arcmin after 20sec maximum S/A damping time– after slow slew - within 1.25 arcmin, (assumed)– pointing knowledge of < 2” maintained throughout maneuvers– 30 minutes maximum observation time

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GN&C: StrategyGN&C: Strategy

• Build on previous work as much as possible: GLAST, EDGE

• Trade between Reaction Wheels and CMG– EDGE already did Reaction Wheel vs RCS– Trade using a CMG and RW combinations

CMG used for the fast slew requirement RW used for every thing else (slow slews, station keeping, dithers)

– Trade using Ball Aerospace Worldview CMG alone

• Use 2 NFOV star trackers to achieve high accuracy pointing knowledge (2”)

• Use 2 WFOV star trackers to maintain orientation during fast slews (1deg/s)– If the NFOV trackers get lost during fast slews, it will take a minute to re-

establish attitude– Coupling the WFOV knowledge with the NFOV can keep the NFOV tracker

from getting lost

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d2,3

cm

PS

d1,1

Fg

PA

i

Solar Incident Angle

qYaw to Nadir Angle

+X (Thrust vector) cp1

-Z

+Y

Notes:1) This diagram used to identify the physical parameters of the spacecraft.

2) Only need to identify one side of symmetric appendages. (ex: only 1 of 2 opposite solar panels identified)

3) d1-y are y dimensions with torque about either the z or x axis depending on direction of attack.d2-x are x dimensions with torque about either the z or y axis depending on direction of attack.d3-z are z dimensions with torque about either the y or x axis depending on direction of attack.

(To central Body)

Disturbance Equations :1 Solar Torque

Ts = Ps A cos(i) (1+q) L

2 Atmospheric Torque Ta = ½ pV2 Cd A L

3 Magnetic TorqueTm = N I A (Bo Ro/R3)(3sin2L+1)1/2 sin(q)

4 Gravity Gradient TorqueTg = 3u / 2R3 | Iz – Ix,y | sin 2q

A1

Appendage 1

Appendage 3

d1,3

d1,2

Appendage 2

cp2 cp3

A2 A3

Appendage 4

cp4

A4

2-25-09

App(n) dn1-y(m)

dn2-x(m)

dn3-z(m)

App Dim(m)

Area(m2)

1 - SA 3.50 0.0 -0.15 3.0 x 2.5 7.50

2 - WFS -0.624 0.55 1.95 1.3 x 1.09 Ø 1.42

3 - WFI 0.620 0.40 1.95 1.3 x 0.80 Ø 1.04

4 - SC 0.0 1.30 -0.05 2.7 x 2.6 Ø 7.02

cp to cm

Ixx = 3317 kgm2

Iyy = 3092 kgm2

Izz = 1900 kgm2

GN&C: ACS Tool InputsGN&C: ACS Tool Inputs

Dimensions are taken form configuration diagram, see configuration charts.

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GN&C: Results and ConclusionsGN&C: Results and Conclusions

• Suggest using Ball Aerospace Worldview Control Moment Gyro 4 wheel set

– One set 4 CMG wheels to perform the fast slews, slow maneuvers, and station keeping– Wheels mounted in a pyramid configuration near the spacecraft center of mass

• Slightly better performance can be achieved using a CMG and Reaction Wheel combination set, but would be higher mass and power, and be significantly more complex• A set of magnetic torquer rods used to perform the de-saturation of the wheels

– Suggest using1 Zarm/Microcosm MT400-2 rods, with .014Nm average torque capability per orbit

De-saturation analysis has not been performed. De-saturation times may be significant, impacting science time Suggest using operational maneuvers to non-GRB event targets in round about paths to de-saturate wheels

• Suggest 2 sets of star trackers– One set of 2 NFOV perpendicular to each other, used for the high accuracy pointing knowledge (2”)

Goodrich has stated that the HD-1003 next generation star tracker can achieve 1” accuracy in x and y A second tracker is needed for the third axis high accuracy knowledge

– Another set of 2 WFOV trackers is suggested for maintaining orientating knowledge during fast slews

AeroAstro Mini-Star Tracker has a 10deg/sec rate capability advertised

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• Suggest using Ball Aerospace M-95 CMG 4 wheel pyramid configuration for all slews, station keeping, and observations.

• Provides up to 6.1 Nm torque(~4.0 Nm required for Xenia)

GN&C: Results GN&C: Results

BallBall Aerospace Worldview CMGAerospace Worldview CMG

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GN&C: ResultsGN&C: Results

Number of

Wheels

Source and Type(All Pyramid Configurations)

Nominal Wheel Condition

1 Wheel Failure Condition Masses Power

Slew Time (min)

Science Times (hr)

Slew Time (min)

Science Times (hr)

Total Wheel

mass (kg)

Total System mass (kg)*

Total System Power (W)

8Ball Aerospace CMG-M95 and Teldix RSI 50-220/45 0.53 3.59 0.79 1.65 167.2 334.4 820

4 Ball Aerospace CMG-M95 0.9 3.22 1.4 0.98 130.8 261.6 2206 Teldix MWI 30-400/37 3 1.34 5 0.79 91.8 183.6 18004 Teldix MWI 30-400/37 5 0.98 15 0.69 61.2 122.4 1200

*Total system mass includes isolation mounts and electronics

Performance Trade Table

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0

1

1

2

2

3

3

4

4

0 1 2 3 4 5 6 7

Scie

nce

Tim

e (h

rs)

60 deg Slew Time (min)

Xenia - CMG vs Reaction Wheel performance tradeScience Time vs Slew Time

(All Pyramid Configurations)

Nominal Wheel Conditions1 Wheel Failure Conditions

4 Teldix-MWI30/400

4 Ball CMG-M95 and 4 Teldix-RSI50/220 Combo Set

4 Ball CMG-M95

6 Teldix-MWI30/400

18 min

2

90 min

1

3

4

5

Shaded area meets reqiurment of 60 deg slew in less then 60 sec

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0

5

10

15

20

25

30

35

40

45

0 5 10 15 20

Nu

mbe

r of

Pos

sibl

e Sl

ews

per

Des

atu

rati

ion

Cyc

le

Equal Observation Times (min)

Ball Aerospace CMG-M95 and Teldix RSI 50-220/45

Ball Aerospace CMG-M95

Teldix MWI 30-400/37

Teldix MWI 30-400/37

Total tob

(min)

Number of slews possible per de-saturation cycle given equal observation times

System Trades

A B C D

17.00 13 11 5 312.00 18 16 7 510.00 22 19 8 68.00 27 24 10 7

5 43 39 16 12

A

B

C

D

Number of slews possible per de-saturation cycle

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AvionicsAvionics

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Avionics: GR&A Avionics: GR&A

• Communication transmission link via TDRSS

• Total science down-link communication data rate: 3.8 Mbps, orbital average

• Total telemetry up-link and down-link communication data rate: 4 kbps per transmission

• Total science on board memory required: 4 Gigabit

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7 –TDRSS at GEO (35888 km)

Supports: Ka-band 27.5-22.2 GHz

25 Mbps up, 800 Mbps dw Ku-band 15.0-13.7 GHz

25 Mbps up, 300 Mbps dwS-band 2.3-2.0 GHz

300 kbps up, 6 Mbps dw

Science data3.8 Mbps orbital average

Ku-band

GRB alerts, TOOEngineering data4 kbps, S-band

Notes:This communication strategy is similar to FERMI (formerly GLAST), and suggested in EDGE.

24 –GPS at GEOL-band 1.575-1.227 GHz

50 bps

WSGTWhite Sands NM

GSFCGoddard MD

GRGTGuam

Xenia600 km orbit

Avionics: Communication StrategyAvionics: Communication Strategy

3 - Ground stations

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Avionics: ResultsAvionics: Results

Astrionics Mass (kg) Power (W)

Attitude Control System 320 240

Command and Data System 22 107

Instrumentation and Monitoring 5 7

Communications System 45 203

Avionics Cabling 34 N/A

Totals 426 557

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Avionics: ResultsAvionics: Results

• An omni directional Ku band communication link was chosen for simplicity and mass savings– A pointing antenna may be blocked by S/C structures, restricting continues

transmission capability– A 4Mbps omni link can be made to TDRSS with a 10w transmitter

A link budget analysis was performed.

• A redundant 5w S-band system is used for command and telemetry links with TDRSS– It is planned to have no direct link to ground for normal operations, all links

are through TDRSS

• The Saab Ericsson Spacecraft Computer has built in redundancy, extra memory and speed capacity, and all the I/O required for this application, along with good heritage

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PowerPower

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Power: GR&APower: GR&A

• Long Mission: 5 Year Desired Life

• 600 km circular orbit: Max Dark Period 35.5 min, Min Light Period 61.2 min.

• Spacecraft must be independently oriented to view events of scientific interest

• Relatively high power levels (1-2 kW) required for science package

• Conditioned power, multiple voltages from common power bus @ 28V

• Required Power: 2027 W (including 30% margin)

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Power: Design HighlightsPower: Design Highlights

• Solar Array – 14.65 m^2– GaAs 3j rated 348 W/m^2 (before Knockdowns)– 2.24 kg / m^2– Inherent Degradation 0.85– Degradation Rate 0.03/yr

• Secondary Batteries – 8 Cells per Unit, 2 Units– Based on Saft Li-Ion VES 180 Cells (50 Ah, 3.6V)– 1.29 Packing Factor– Cell Load Balancing Electronics– Max Depth of Discharge < 40%

• Array Regulation – Direct Energy Transfer (0.95 Efficiency)

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Power Masses Qty 169.52 kg

PDU 1 12.48 kg 12.48 kg

5 m, redundant Cabling 1 5.59 kg 5.59 kg

ARU 1 31.35 kg 31.35 kg

Solar Array 1 32.82 kg 32.82 kg

2880 Wh Secondary Battery 2 11.66 kg 23.32 kg

Battery Charger 1 63.97 kg 63.97 kg

Power: ResultsPower: Results

Sized to 2027W End of Life Power (1758W after 10 Years)

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ThermalThermal

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Thermal Control: GR&AThermal Control: GR&A

• Primary objective is to develop a passive thermal design concept for the Xenia spacecraft. Heat rejection of instrument and subsystems power is accomplished by spacecraft radiators, closeout MLI, heat pipes and silverized Teflon tape.– Circular orbit, altitude 600 km

– 5° inclination, βmax = 33.5 °, βmin = 0 °– 3-axis stabilized, 45° sun avoidance angle

• Spacecraft bus outer structural panels double as radiators– Spacecraft Bus composed of Aluminum plate (thickness varies) for optimal

thermal conductivity.– Radiator panels located on the sides and bottom of the spacecraft

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Thermal Control: DesignThermal Control: Design

Side Panel Radiator/ 2 plcsSilverized Teflon Tapeα/ε=.2/.7 (Degraded)

Bottom Panel Radiatorα/ε=.2/.7 (Degraded)

Interior surfacesBare Aluminum

Closeout Blankets, 2 layerAluminized Teflonα/ε=.14/.62

The radiator panels on the ISS and Shuttle are covered with silver coated FEP tape. To insure a long life in the presence of atomic oxygen, the tapes were coated with silicon oxide which acts as an atomic oxygen absorber.

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Thermal Control: AnalysisThermal Control: Analysis

WFI

WFMCMGs

ChargerFlight Comp.DAUIMU

Battery

CMGs

WFS Electronics

Heat Pipesused to isothermalize heat loads

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Thermal Control: AnalysisThermal Control: AnalysisComponent Total Power(w)

Science InstrumentsWFS 909.00Detector head 0.00FEE 23.00Filter wheel 10.00Mirror and casing 150.00Digital electronic box 135.00Control/Power Electronic box 103.00Crysyatic cooler 0.002ST Drive Electronics box 213.002ST Drive Electronics box 399.00ADR Analog Control box 59.00WFI 60.00Camera head and FEE 30.00Filter wheel and shutter 20.00TEC 35.00Mirror and casing 40.00Analog Electronic box 25.00Control/Power Electronic box 35.00WFM 92.00Detector 48.00ICU Electronic box 92.00

Component Total Power(w)ACS (Attitude Control Sytem) 140.00Sun Sensor Electronics Unit 6.00

Star Tracker WFOV 4.00

Star Tracker NFOV 20.00

Magnetometers - 3 axis each 4.00

IMU Assy 20.00CMG Controllers (4) 60.00Control Moment Gyro 60.00Magnetic torqueres 10.5

CDS (Command and Data System) 82.00Flight Computer 60.00

Data Aquistion Unit 22.00

Communications 156.00MGA Transmitter/Amplifier 114.00LGA Transceiver/Transponder 36.00GPS units w/combiner 6.00Power and ThermalPower System 585.00ARU 109.00Secondary Battery (2) 104.00Battery Charger 364.0028 VDC PDU 8.00Thermal System (Heaters) 15.00

Total Spacecraft Heat Dissipation considered in analysis = 1776W

49

Thermal Control: ResultsThermal Control: Results

• Spacecraft temperatures are -10° to 34° C • Heat dissipation = 1776W • β=33.5°• Sun Angle = 90°• Orbital Average Temperatures (°C)

50


• Spacecraft temperatures are -9° to 35° C • Heat dissipation = 1776W • β=0°• Sun Angle = 90°• Orbital Average Temperatures (°C)

51


Component Units Mass (kg)

Instrument Light Shield & Baffle MLI

15 m2 5.0

Closeout blankets 15 m2 7.5

Heat Pipes 4 @ 1.3 kg each 5.2

Silverized Teflon Tape

25@ .6 kg/m2 15.0

Total 32.7

52

PropulsionPropulsion

53

Propulsion: GR&APropulsion: GR&A

• Spacecraft initial mass – 2760 kg• Deorbit V 163 m/sec (from mission analysis)• T/W > 0.025• Engine

– 1 Aerojet R42 Engine Oxidizer – NTO Fuel – MMH Isp = 303 seconds 5% residual

• Tank configuration– 1 tank for each propellant

Metallic tanks Pressure = 240 psia

– Separate pressurization system for each propellant 2 helium bottles Initial pressure at 4500 psia

54

Propulsion: System SchematicPropulsion: System Schematic

GHe Tank (each)V = 0.005 m3 (326 in3)Geometry = 0.22 m sphereMEOP = 4,500 psia

Prop Tanks (each)V = 0.056 m3 (3418 in3)Geometry = 0.62m X 0.20 m dia.MEOP = 240 psia

He

Ox

He

Fu

55

Propulsion: Mass StatementPropulsion: Mass Statement

WBS Element - Descent Stage Qty Unit Mass (kg) Total Mass (kg)2.0 15.50

2.1 Main Engines 1 5.00 5.002.2 Fuel Tank 1 3.30 3.302.3 Main Oxidizer Tank 1 3.60 3.602.4 Pressurization Tank 2 1.30 2.602.5 Feed System 1 1.00 1.00

8.0 4.658.3 30% 4.65

20.159.0 6.10

9.1 Propellant Residuals 1 5.749.1.1 Fuel 1 2.17 2.179.1.2 Oxidizer 1 3.57 3.57

9.2 Pressurant 0.369.2.1 Fuel 1 0.18 0.189.2.2 Oxidizer 1 0.18 0.18

6.10Total Less Propellant 26.2512.0 137.26

12.1 1 51.83 51.8312.2 Main Oxidizer 1 85.43 85.43

Gross Mass 163.51

Main Fuel

Growth

Dry Mass

Propellant

Inert Mass

Non-Cargo

Propulsion

Propulsion

56

StructuresStructures

57

Structures: GR&AStructures: GR&A

• Spacecraft Bus– Aluminum 2024-T351 plate for durability and optimal thermal conductivity– 7075-T651 Al used for struts and adapter ring– Two exterior structural panels and aft aluminum panel double as radiators– Half of the outer surface panel area is required for thermal management– The rest of the exterior is closed out with Multi-Layer Insulation (MLI)

• Secondary Structure – WFI and WFS payload mass is distributed along the axis of each

instrument and supported with secondary structure– A folding boom and vibration damping mechanism to minimize oscillations

after fast slew is similar to one used on the Hubble telescope, but proportionally less massive

58

Structures: Loads GR&AStructures: Loads GR&A

• Maximum Launch Loads for Falcon 9 payload– 5.0 g along launch axis– 0.9 g lateral to launch axis

• Load Set– Axial plus lateral at 45 degree intervals around bus

• Strength Criteria– Factor of Safety (FOS) 1.4– Positive Margin of Safety (MOS) for static launch load analysis– No buckling or stiffness analysis performed

• Optimization– 3/5 of the structure could be composite– Structures mass is expected to decrease when analysis optimization is

complete

59

Structures: AnalysisStructures: Analysis

• Finite Element Modeling and Post-processing (FEMAP)– Images showing structural analysis results

60

Structures: Results

Structural Mass Qty Total 489* kg

Secondary

Solar Panel Structure Solar array dampers, actuators, and booms 2 30.5 61.0

Secondary Propulsion 1 15 .0 15 .0

Secondary Science Instruments 1 100.0* 100.0*

Primary Spacecraft Bus 1 313.0* 313.0*

* Further optimization and analysis could decrease the structural mass of these components.

61

ConclusionsConclusions

62

ConclusionsConclusions

• Observatory fits within the Falcon 9 mass and volume envelope– Plenty of payload margin when launching from Omelek.

• Pointing, slow slewing, and fast slewing requirements met– The use of control moment gyros (CMGs) enables the observatory to meet

these opposing requirements, even with one wheel failure.

• Thermal requirements met– Thermal analysis of the Xenia spacecraft with all instrument electronic units

and subsystem heat loads considered, resulted in an internal temperature range of -10C to 35C. This temperature range is well within the operating temperature range of all instruments and subsystem components located within the spacecraft.

63

BackupBackup

64

Mission Analysis: Orbital Lifetime Mission Analysis: Orbital Lifetime InputsInputs

Model Parameters

Atmosphere NRL MSISE 2000

Solar flux sigma value 2

Rotating Atmosphere no makes calculations more conservative

Satellite Parameters

Drag coefficient 2.2

Drag area 24 m2

Area exposed to sun 30 m2

reflection coefficient 1

Satellite mass 2100 – 2500 kg

65

Mission Analysis: GR&A for Falcon9Mission Analysis: GR&A for Falcon9

• Performance computations are based on the following main assumptions:

– This performance does not include the effects of orbital debris compliance, which must be evaluated on a mission specific basis. This could result in a significant performance impact for missions in which launch vehicle hardware remains in Earth orbit

– This performance is reflective of the Block 2 version of the Falcon 9

– 3-sigma mission required margin, plus additional reserves determined by the LSP

– A payload adapter has been assumed

• Source: NASA LSP performance quote.

66


Required Delta-V for Deorbiting Satellite from Circular Orbit for

Various Interface Flight Path Angles

-2-1.8

-1.6-1.4

-1.2-1

60

80

100

120

140

160

180

200

220

200 300 400 500 600 700Initial Circular Altitude (km)

Delta-V (m/s)

flight path angle at interface (deg)

interface defined as 400000ft altitude

67


Perigee Altitude for Deorbiting Satellite from Circular Orbit for

Various Interface Flight Path Angles

-2-1.8-1.6

-1.4-1.2-1

-400

-200

0

200

400

600

800

1000

1200

1400

200 300 400 500 600 700

Initial Circular Altitude (km)

Perigee Altitude (km)

flight path angle at interface (deg)

interface defined as 400000ft altitude

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Deorbit Gravity Loss vs. Thrust-to-Weight Ratiofor 600km Circ, Isp=250s

(Add this to the 161.3 m/s Impulsive Delta-V)

0

1

2

3

4

5

6

7

8

9

10

0.01 0.1 1 10Thrust-to-Weight Ratio

Gravity Loss (

m/s)

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Xenia - Spacecraft Avionics Functional Diagram

GNC/Vetronics/CommFlight Computer

Systemwith internal dual redundancy

De-orbitEngine

Controller

Data Acquisition

Unit

Igniters

S-bandTransceiver

2 – NFoV Star Tractor

2 – WFoV Star Tractor

2 - IMU Gyros

3 - Magnetometer

4 - Sun Sensors

S-Band4 kbps

Sensors

Amplifier

Ku-Band4 Mbps

Transponder

28 VDCPower

DistributionUnit

ArrayRegulator

Unit

Solar Array #1

Charger

Communications System (3)

Attitude Control System (ACS) (4)

Power System (2)

Survival Heaters

Command and Data System (CDS) (1)

TorqueMotors

GimbleMotors

Solar Array #2

MagneticTorque Rods

WFI WFS

KaTransceiver

Science Instrument Bus (1) (2)

Heaters

H&M Instrumentation-Pressure

-Temperature

-Strain

Deploymentand

PointingMotors

2-23-09

AnalogElectronic

UnitDigital Unit

WFM

HUPS 1

HUPS 2

Controller and PowerElectronics Unit

ControllerAnd

Power Unit

ICUElectronicController

Comm. ToGroundTDRSS

GPS

6

4 typ 2 typ

vv3-Magnetic

Torquer set

Includes2-16 Gbits

Memoryboards

vv

v4-CMGwheels

PowerAmp

2ST Drive

ADR Analog ST/JT Drive

SecondaryBattery

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Avionics Avionics -- SystemSystem

Functional Diagram Notes1. It is assumed that the individual science instrument packages include all the required data

processing, filtering, and buffering required, along with thermal, health, and status control. All science data is to be transmitted to the spacecraft computer via a dual redundant spacecraft data bus for storage and downloading to ground. Instrument health and status telemetry will be collected by a second dual redundant spacecraft data bus, processed, and stored independently of the science data. All science and telemetry data should be ID and time stamp for later correlation and downloading to ground.

2. A dual redundant primary power feed will be supplied to an instrument controller for each of the 4 major instruments. Those controllers must distribute secondary power to the instrument and instrument's electronic boxes, perform all required operations (e.g. safe mode), control any mechanism required (e.g. shutters), and perform the thermal management of the dedicated systems. All cabling between the controllers and science boxes should be included in the science package mass estimates.

3. The communication system will be similar to GLAST/FERMI. The TDRSS communication satellite system will be used as the primary means of uploading commands and targets of opportunity (TOO), downloading science and telemetry data, along with broadcasting a detected event. Since the intent is to keep the comm system small and simple, direct ground link will be used only as backup at low data rates.

4. The final study plan is to perform fast slews with a 4 wheel Control Moment Gyro (CMG) set provided by Ball Aerospace. The slow target slews and regular station keeping operations will also be done using the same CMG set. De-saturation of the wheels will be accomplished using electromagnetic torque rods when needed. De-saturation down time may be significant unless de-saturation maneuvers are done.

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ACS Tool Outputs and analysisACS Tool Outputs and analysis

3,317 3,092 1,900

Note: Desaturation period = 1 orbit (~90 min) Units Ixx

kgm2 (roll) Iyy

kgm2 (pitch) Izz

kgm2 (yall)

Disturbance Torques

Solar Torques Nm 8.195E-06 1.245E-06 8.958E-07

Atmospheric Torques Nm 1.707E-03 1.153E-03 1.703E-04

Gravity Torques Nm 2.604E-03 2.190E-03 4.134E-04

Total disturbance Torques Nm 4.319E-03 3.345E-03 5.847E-04

Disturbance times (.75' drift) sec 12.94 14.20 26.63

Disturbance times (1.25' drift) sec 16.71 18.33 34.37

Attitude Corrections

Correction Torque (.75' drift) Nm 0.0417 0.0389 0.0239

Correction Torque in (1.25' drift) Nm 0.0250 0.0233 0.0143

Correction time (.75' drift) sec 8.333 8.333 8.333

Correction time (1.25' drift) sec 13.88 13.88 13.88

Correction Momentum (both drifts) Nms 0.174 0.162 0.099

Correction Cycles per desat period (.75' drift) # 254 240 155

Correction Cycles per desat period (1.25' drift) # 177 168 112

Momentum per desat period (.75' drift) Nms 44.11 38.86 15.42

Momentum per desat period (1.25' drift) Nms 30.74 27.20 11.14

Maneuvers per desaturation Period

1 - Fast Slew Torque (60deg/45sec) Nm 6.861 6.396 3.930



1 - Slow Slew Torque (100deg/100sec) Nm 2.316 2.159 1.326

1 - Fast Slew Momentum (60deg/45sec) Nms 154.38 143.91 88.43



1 - Slow Slew Momentum (100deg/75sec) Nms 154.38 143.91 88.43



Sum of greatest Momentums Nms 198.49 182.76 103.85

Sum of mid-level Momentums Nms 159.90 146.79 81.74

Sum of least Momentums Nms 121.30 110.81 59.64

Critical design parameters

Recommended design parameters

Ball Aerospace M95 CMG - 129 Nms each wheel x 2.31 for a 4 wheel pyramid = 298 Nms.

Collective torque capability = 6.1 Nm

72

ESA planned and heritage missions:• Herschel - 3.5 m IR telescope at L2• Planck - cosmic microwave background• Pleiades - Earth observation satellite • Aeolus - Atmospheric wind sensor(same as MSFC canceled Sparkle)

Saab Ericsson Space (Sweden) – Spacecraft Management Computer

AOCS Interfaces• GPS Receiver• Magnetometer interfaces• Gyro interfaces• Reaction wheel interfaces• Star tracker interfaces• Thruster control interfaces• Sun sensor interfaces

Properties and Interfaces• Power consumption: <40 W average, < 60 W peak • Mass: 18 kg• Dimensions: 420 (L) x 270 (H) x 276 (D) mm• Reliability:>0,99 over a 3-year mission using class B components>0,95 over a 15-year mission using class S Components

• Heaters: 50 W per line, > 500 W total• secondary power distribution• Solar Array Drive Motor• 32/64 Gbit mass memory boards• 1553 Data buses• 40 Mbps payload wire links• 20 Mbps RS-422 Synchronous serial links• 1.5 Mbaud UART links RS-422 or RS-485• RS-422 Synch pulses fixed and programmable

Avionics Avionics –– Flight ComputerFlight Computer

73

Avionics Avionics -- CommunicationsCommunications

74

Avionics Avionics -- CommunicationsCommunications

75

GN&G GN&G –– Attitude SensorsAttitude Sensors

76

GN&G GN&G –– Attitude SensorsAttitude Sensors

77

•Suggest using Ball AerospaceM-95 CMG 4 wheel pyramid configuration for all slews, station keeping, and observations.

•Provides up to 6.1 Nm torque(4.8 Nm required for Xenia)

BallBall Aerospace Worldview CMGAerospace Worldview CMG

GN&G GN&G –– CMGCMG

78

•Option using MWA-50 wheels for slow slewstation keeping, and dither maneuvers.

•Provides 68Nms momentum storageand .07Nm torque

93 to 58 Nms required for mid-level performance.05Nm torque required for dither roll

•4 wheels at 10.5kg each gives a total mass of 42kg

GN&C GN&C –– Reaction Wheel example Reaction Wheel example

79

Power: PhotoPower: Photo--Voltaic ArraysVoltaic Arrays

• Ga-As 3j• 258 W / m^2 (End of Life, with knockdowns for cell

mismatch, interconnect failures, margin)• 2.24 kg / m^2

80

Power: Array Regulation UnitPower: Array Regulation Unit

• Sequential Shunt Regulator• 60 Strings• PWM Freq 50kHZ• Ripple < 1.65 %

81

Power: Battery Charge/Discharge Power: Battery Charge/Discharge Units (BCDU)Units (BCDU)

• Linear Regulation• Charge / Discharge Efficiency 81%• Ripple < 0.5 %

82

Thermal: Heat Pipe InfoThermal: Heat Pipe Info

Ambient temperature heat pipes have been used successfully in numerous spacecraft applications. They are accepted as a reliable aerospace component based on extensive flight data. One of the most extensive application in the use of heat pipes aboard an operational spacecraft has been on the Applications Technology Satellite (STS-6). A total of 55 heat pipes were placed in equipment panels to carry solar and internal power loads to radiator surfaces. Ammonia was used with aluminum axially grooved tubing. Data taken over a 24 hour orbital period shows a maximum gradient of 3 deg. C existed from one side of the spacecraft to the other. No degradation in thermal design was seen.

Axially grooved aluminum extrusions with ammonia working fluid are used to isothermalize the equipment platform of the International Ultraviolet Explorer as shown in the figure below

83

Propulsion: RPropulsion: R--42 Engine42 EngineData SheetData Sheet

R-42 890N (200 lbf) BIPROPELLANT ROCKET ENGINE

Design Characteristics• Propellant . . . . . . . . . . . . . . . . . . . . . . . . . . . . MMH/NTO(MON-3)

• Thrust/Steady State. . . . . . . . . . . . . . . . . . . . . . . . . . . 890N (200 lbf )

• Inlet Pressure Range . . . . . . . . . . . . . . . 29.3-6.9 bar (425-100 psia)

• Chamber Pressure*. . . . . . . . . . . . . . . . . . . . . . . . 7.1 bar (103 psia)

• Expansion Ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160:1

• Flowrate*.. . . . . . . . . . . . . . . . . . . . . . . . . . .300 g/sec (0.66 lbm/sec)

• Valve . . . . . . . . . . . . . . . Aerojet Solenoid, Single Coil, Single Seat

• Valve Power . . . . . . . . . . . . . . . . . . . . . . . . . . . 46 Watts @ 28 Vdc

• Mass. . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . .4.53 kg (10.0 lbm)

*At rated thrust

Performance• Specif ic Impulse* . . . . . .. . . . . . . . . . . . 303 sec (lbf-sec/lbm)

• Total Impulse . . . . . . . 24,271,000 N-sec (5,456,700 lbf-sec)

• Total Pulses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134

• Minimum Impulse Bit . . . . . . . . . . 44.48 N-sec (10.0 lbf-sec)

• Steady State Firing Cumulative . . . . . . . . . . . . . . 27,000 sec

• Steady State Firing (Single Firing) . . . . . . . . . . . . . 3,940 sec

15.34

25.006.0

Dimensions are in inches

Approved for public release and export

Rev. Date: 5/17/06

11411 139th PL NE • P.O. BOX 97009 • REDMOND, WA 98073-9709(425) 885-5000 FAX (425) 882-5747

Reference• AIAA - 1990 - 2055

Documents

Spacecraft Design March 2, 2009